This invention relates to turbine engine combustor air-cooled liner segments and, more particularly, to repair of a damaged edge portion of such segment.
One form of a turbine engine, for example an axial flow gas turbine engine to propel aircraft or marine vessels or to generate electrical power, includes a combustion section disposed generally between an axially forward compressor section and an axially aft turbine section. As is well known in the gas turbine engine art, air from the compressor section is mixed with fuel in the combustion section and ignited to provide hot expanding products of combustion for extraction of power by the turbine section. Such combustion of fuel in a rapidly flowing pressurized air atmosphere generates very strenuous high temperature environmental oxidizing and corrosive conditions along with highly erosive conditions all of which can damage components of the combustion section.
One form of a currently used combustion section is called an annular combustor. As used herein, terms such as “annular”, “radial”, “circumferential”, “axial”, etc. refer to directions about a typical axial flow gas turbine engine. One form of an annular combustor comprises an outer annular frame-like member carrying within its annular interior in which combustion is conducted at least one fuel nozzle, baffles, and a combustor liner, for example comprising a plurality of generally arcuate combustor liner segments. Such segments, typically precision cast from a high temperature alloy based on at least one of Fe, Co, and Ni interface between and protect the outer frame from conditions within the combustor as well as guide the combustion and its products. In one embodiment, each such combustor liner segment is protected on its radially inner surface with a commercially available thermal barrier coating (TBC), one example of which primarily is yttria stabilized with zirconia. In addition, cooling air is passed over the segments' radially outer surfaces that include there-across a plurality of spaced-apart generally radially outward extending protuberances or pin-like structures functioning as heat exchange surfaces or members and designed with spaces therebetween to control cooling air flow about the protuberances.
In their annular, circumferentially arcuate disposition about the combustor interior to define the combustor liner, the segments partially overlap one another axially downstream so that the cooling air traversing the radially outer surface of a segment between the protuberances is discharged over a portion of the TBC of a superimposed adjacent segment. Nevertheless, it has been observed after service operation, that the strenuous engine operating conditions can result in damage to or erosion of a downstream edge portion of certain combustor liner segments. Because such cast segments, including the radially outer spaced-apart pin-like structure, is relatively expensive to manufacture, typically by a lost wax type precision casting, it is desirable to repair rather replace such a member. However, a low cost repair method has not been available and damaged liner segments have been discarded.
Repair by bonding of a replacement edge portion by typical current bonding methods can result in flow of excessive metal flow about such protuberances. For example, current brazing methods including disposing brazing alloy at a face or surface that includes the protuberances, or bonding by typical fusion welding that melts at least a portion of parent metal as well as any weld metal, can result in excessive brazing or molten alloy flow. Such excessive flow can block, interfere with, and/or change a designed pattern and/or amount of cooling airflow on the segment radially outer surface. Provision of a segment repair method that maintains the integrity of the radially outer surface designed cooling air-flow control and, in one preferred form provides the segment with enhanced oxidation resistance at an operating temperature, can improve and enable repair rather than replacement of damaged combustor liner segments.